Gas turbine engine airfoil

ABSTRACT

A gas turbine engine includes a combustor section arranged between a compressor section and a turbine section. The compressor section includes at least a low pressure compressor and a high pressure compressor. The high pressure compressor is arranged upstream of the combustor section. A fan section is included. The low pressure compressor is downstream from the fan section. An airfoil is arranged in the low pressure compressor and includes pressure and suction sides extending in a radial direction from a 0% span position to a 100% span position. The airfoil has an axial stacking offset at a 10% span position that is positive and greater than or equal to an axial stacking offset at the 100% span position. The axial stacking offsets measured relative to an aftward direction.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. application Ser. No.16/744,749 filed Jan. 16, 2020, which is a continuation of U.S.application Ser. No. 15/115,360 filed Jul. 29, 2016, which is a 371 ofInternational Application No. PCT/US2015/016018 filed Feb. 16, 2015,which claims priority to U.S. Provisional Application No. 61/941,685,which was filed on Feb. 19, 2014 and is incorporated herein byreference.

BACKGROUND

This disclosure relates to gas turbine engine airfoils. Moreparticularly, the disclosure relates to airfoil axial stacking offsetin, for example, a gas turbine engine compressor.

A turbine engine such as a gas turbine engine typically includes a fansection, a compressor section, a combustor section and a turbinesection. Air entering the compressor section is compressed and deliveredinto the combustor section where it is mixed with fuel and ignited togenerate a high-speed exhaust gas flow. The high-speed exhaust gas flowexpands through the turbine section to drive the compressor and the fansection. The compressor section typically includes at least low and highpressure compressors, and the turbine section includes at least low andhigh pressure turbines.

Direct drive gas turbine engines include a fan section that is drivendirectly by one of the turbine shafts. Rotor blades in the fan sectionand a low pressure compressor of the compressor section of direct driveengines rotate in the same direction.

Gas turbine engines have been proposed in which a geared architecture isarranged between the fan section and at least some turbines in theturbine section. The geared architecture enables the associatedcompressor of the compressor section to be driven at much higherrotational speeds, improving overall efficiency of the engine. Thepropulsive efficiency of a gas turbine engine depends on many differentfactors, such as the design of the engine and the resulting performancedebits on the fan that propels the engine and the compressor sectiondownstream from the fan. Physical interaction between the fan and theair causes downstream turbulence and further losses. Although some basicprinciples behind such losses are understood, identifying and changingappropriate design factors to reduce such losses for a given enginearchitecture has proven to be a complex and elusive task.

Prior compressor airfoil geometries may not be suitable for thecompressor section of gas turbine engines using a geared architecture,since the significantly different speeds of the compressor changes thedesired aerodynamics of the airfoils within the compressor section.Counter-rotating fan and compressor blades, which may be used in gearedarchitecture engines, also present design challenges.

SUMMARY

In one exemplary embodiment, a compressor airfoil of a turbine enginehaving a geared architecture includes pressure and suction sides thatextend in a radial direction from a 0% span position to a 100% spanposition. The airfoil has a relationship between an axial stackingoffset and span position that includes a curve with a negative slopefrom 90% span to 100% span. The negative slope leans forward relative toan engine axis.

In a further embodiment of the above, the curve has a negative slopefrom 80% span to 100% span.

In a further embodiment of any of the above, the curve has a negativeslope from 70% span to 100% span.

In a further embodiment of any of the above, the curve has a positiveslope beginning at 0% span. The positive slope leans aftward relative tothe engine axis.

In a further embodiment of any of the above, the positive slope extendsfrom 0% span to 40% span.

In a further embodiment of any of the above, the curve transitions fromthe positive slope to the negative slope in a range of 40% span to 75%span.

In another exemplary embodiment, a gas turbine engine includes acombustor section that is arranged between a compressor section and aturbine section. A fan section has an array of twenty-six or fewer fanblades. The fan section has a fan pressure ratio of less than 1.55. Ageared architecture couples the fan section to the turbine section orthe compressor section. An airfoil is arranged in the compressor sectionand includes pressure and suction sides that extend in a radialdirection from a 0% span position to a 100% span position. The airfoilhas a relationship between an axial stacking offset and span positionthat includes a curve with a negative slope from 90% span to 100% span.The negative slope leans forward relative to an engine axis.

In a further embodiment of any of the above, the compressor sectionincludes at least a low pressure compressor and a high pressurecompressor. The high pressure compressor is arranged immediatelyupstream of the combustor section.

In a further embodiment of any of the above, the airfoil is provided ina compressor outside the high pressure compressor.

In a further embodiment of any of the above, the low pressure compressoris counter-rotating relative to the fan blades.

In a further embodiment of any of the above, the gas turbine engine is atwo-spool configuration.

In a further embodiment of any of the above, the low pressure compressoris immediately downstream from the fan section.

In a further embodiment of any of the above, the airfoil is rotatablerelative to an engine static structure.

In a further embodiment of any of the above, the curve has a negativeslope from 80% span to 100% span.

In a further embodiment of any of the above, the curve has a negativeslope from 70% span to 100% span.

In a further embodiment of any of the above, the curve has a positiveslope beginning at 0% span. The positive slope leans aftward relative tothe engine axis.

In a further embodiment of any of the above, the positive slope extendsfrom 0% span to 40% span.

In a further embodiment of any of the above, the curve transitions fromthe positive slope to the negative slope in a range of 40% span to 75%span.

BRIEF DESCRIPTION OF THE DRAWINGS

The disclosure can be further understood by reference to the followingdetailed description when considered in connection with the accompanyingdrawings wherein:

FIG. 1 schematically illustrates a gas turbine engine embodiment with ageared architecture.

FIG. 2 schematically illustrates a low pressure compressor section ofthe gas turbine engine of FIG. 1.

FIG. 3 is a schematic view of airfoil span positions.

FIG. 4 is a schematic view of a cross-section of an airfoil sectioned ata particular span position and depicting directional indicators.

FIG. 5 graphically depicts curves for several example airfoil axialstacking offset relative to span, including two prior art curves andseveral inventive curves according to this disclosure.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmenter section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core flow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures. That is, the disclosedairfoils may be used for engine configurations such as, for example,direct fan drives, or two- or three-spool engines with a speed changemechanism coupling the fan with a compressor or a turbine sections.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis X relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisX which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five (5:1). Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicyclic geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

The example gas turbine engine includes the fan 42 that comprises in onenon-limiting embodiment less than about twenty-six (26) fan blades. Inanother non-limiting embodiment, the fan section 22 includes less thanabout twenty (20) fan blades. Moreover, in one disclosed embodiment thelow pressure turbine 46 includes no more than about six (6) turbinerotors schematically indicated at 34. In another non-limiting exampleembodiment the low pressure turbine 46 includes about three (3) turbinerotors. A ratio between the number of fan blades 42 and the number oflow pressure turbine rotors is between about 3.3 and about 8.6. Theexample low pressure turbine 46 provides the driving power to rotate thefan section 22 and therefore the relationship between the number ofturbine rotors 34 in the low pressure turbine 46 and the number ofblades 42 in the fan section 22 disclose an example gas turbine engine20 with increased power transfer efficiency.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.55. In another non-limiting embodimentthe low fan pressure ratio is less than about 1.45. In anothernon-limiting embodiment the low fan pressure ratio is from 1.1 to 1.45.“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram °R)/(518.7° R)]°^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1200ft/second (365.7 meters/second).

Referring to FIG. 2, which schematically illustrates an example lowpressure compressor (LPC) 44, a variable inlet guide vane (IGV) isarranged downstream from a fan exit stator (FES). The figure is highlyschematic, and the geometry and orientation of various features may beother than shown. An actuator driven by a controller actuates the IGVabout their respective axes. Multiple airfoils are arranged downstreamfrom the IGV. The airfoils include alternating stages of rotors (ROTOR1,ROTOR2, ROTOR3, ROTOR4) and stators (STATOR1, STATOR2, STATOR3,STATOR4). In the example shown in FIG. 2, the LPC includes four rotorsalternating with four stators. However, in another example, a differentnumber of rotors and a different number of stators may be used.Moreover, the IGV and stator stages may all be variable, fixed or acombination thereof.

The disclosed airfoils may be used in a low pressure compressor of a twospool engine or in portions of other compressor configurations, such aslow, intermediate and/or high pressure areas of a three spool engine.However, it should be understood that any of the disclosed airfoils maybe used for blades or vanes, and in any of the compressor section,turbine section and fan section.

Referring to FIG. 3, span positions on an airfoil 64 are schematicallyillustrated from 0% to 100% in 10% increments. Each section at a givenspan position is provided by a conical cut that corresponds to the shapeof the core flow path, as shown by the large dashed lines. In the caseof an airfoil with an integral platform, the 0% span positioncorresponds to the radially innermost location where the airfoil meetsthe fillet joining the airfoil to the inner platform. In the case of anairfoil without an integral platform, the 0% span position correspondsto the radially innermost location where the discrete platform meets theexterior surface of the airfoil. For airfoils having no outer platform,such as blades, the 100% span position corresponds to the tip 66. Forairfoils having no platform at the inner diameter, such as cantileveredstators, the 0% span position corresponds to the inner diameter locationof the airfoil. For stators, the 100% span position corresponds to theoutermost location where the airfoil meets the fillet joining theairfoil to the outer platform.

Airfoils in each stage of the LPC are specifically designed radiallyfrom an inner airfoil location (0% span) to an outer airfoil location(100% span) and along circumferentially opposite pressure and suctionsides 72, 74 extending in chord between a leading and trailing edges 68,70 (see FIG. 4). Each airfoil is specifically twisted with acorresponding stagger angle and bent with specific sweep and/or dihedralangles along the airfoil. Airfoil geometric shapes, stacking offsets,chord profiles, stagger angles, sweep and dihedral angles, among otherassociated features, are incorporated individually or collectively toimprove characteristics such as aerodynamic efficiency, structuralintegrity, and vibration mitigation, for example, in a gas turbineengine with a geared architecture in view of the higher LPC rotationalspeeds.

The airfoil 64 has an exterior surface 76 providing a contour thatextends from a leading edge 68 generally aftward in a chord-wisedirection H to a trailing edge 70, as shown in FIG. 4. Pressure andsuction sides 72, 74 join one another at the leading and trailing edges68, 70 and are spaced apart from one another in an airfoil thicknessdirection T. An array of airfoils 64 are positioned about the axis X(corresponding to an X direction) in a circumferential or tangentialdirection Y. Any suitable number of airfoils may be used for aparticular stage in a given engine application.

An axial stacking offset X_(d) corresponds to the location of the centerof gravity X_(CG) for a particular section at a given span locationrelative to a reference point 80 in the X direction. The reference point80 is a location such as the axial center of the root or the axialcenter of a rotor bore, for example. The value X_(d) corresponds to theaxial distance from the reference point 80 to the center of gravity. Apositive X is on the aft side of reference point 80, and a negative X ison the forward side of the reference point 80. A positive slope to theaxial stacking offset profile is where the span section leans aftward,and a negative slope to the axial stacking offset profile is where thespan section leans forward relative to the engine's axis X.

The exterior surface 76 of the airfoil 64 generates lift based upon itsgeometry and directs flow along the core flow path C. The airfoil 64 maybe constructed from a composite material, or an aluminum alloy ortitanium alloy, or a combination of one or more of these.Abrasion-resistant coatings or other protective coatings may be appliedto the airfoil. The rotor stages may be constructed as an integrallybladed rotor, if desired, or discrete blades having roots secured withincorresponding rotor slots of a disc. The stators may be provided byindividual vanes, clusters of vanes, or a full ring of vanes.

Airfoil geometries can be described with respect to various parametersprovided. The disclosed graph(s) illustrate the relationships betweenthe referenced parameters within 10% of the desired values, whichcorrespond to a hot aerodynamic design point for the airfoil. In anotherexample, the referenced parameters are within 5% of the desired values,and in another example, the reference parameters are within 2% of thedesired values. It should be understood that the airfoils may beoriented differently than depicted, depending on the rotationaldirection of the blades. The signs (positive or negative) used, if any,in the graphs of this disclosure are controlling and the drawings shouldthen be understood as a schematic representation of one example airfoilif inconsistent with the graphs. The signs in this disclosure, includingany graphs, comply with the “right hand rule.” The axial stacking offsetvaries with position along the span, and varies between a hot, runningcondition and a cold, static (“on the bench”) condition.

The geared architecture 48 of the disclosed example permits the fan 42to be driven by the low pressure turbine 46 through the low speed spool30 at a lower angular speed than the low pressure turbine 46, whichenables the LPC 44 to rotate at higher, more useful speeds. The axialstacking offset in a hot, running condition along the span of theairfoils 64 provides necessary compressor operation in cruise at thehigher speeds enabled by the geared architecture 48, to thereby enhanceaerodynamic functionality and thermal efficiency. As used herein, thehot, running condition is the condition during cruise of the gas turbineengine 20. For example, the axial stacking offset in the hot, runningcondition can be determined in a known manner using numerical analysis,such as finite element analysis.

FIG. 5 illustrates the relationship between the axial stacking offsetand the average span (AVERAGE SPAN %), which is the average of theradial position at the leading and trailing edges 68, 70. The axialstacking offset assumes a center of gravity based upon a homogeneousmaterial throughout the airfoil cross-section. In one example, theairfoils are LPC rotor blades. Two prior art curves (“PRIOR ART”) areillustrated as well as several example inventive curves 88, 90, 92, 94,96. The airfoil 64 has a relationship between an axial stacking offsetand span position that includes a curve with a negative slope from 90%span to 100% span. The prior art airfoil curves are substantially linearacross the entire span of the airfoil with substantially no offset. Thecurves 88, 90, 92, 94, 96 have a negative slope from 80% span to 100%span in the example. The curves may have a negative slope from at least70% span to 100% span, for example.

The inventive curves 90, 92, 94, 96 have a positive slope beginning at0% span and may extend from 0% span to at least 40% span, for example.The curves transition from the positive slope to the negative slope in arange of 40% span to 75% span, providing a maximum aftward leaning axialstacking offset in this range.

The prior art has generally used straight, axially non-stacked LPCblades. The disclosed airfoils include significant forward and aftwardstacking to improve the aerodynamic efficiency of the high speed LPCblades downstream from a counter-rotating fan.

It should also be understood that although a particular componentarrangement is disclosed in the illustrated embodiment, otherarrangements will benefit herefrom. Although particular step sequencesare shown, described, and claimed, it should be understood that stepsmay be performed in any order, separated or combined unless otherwiseindicated and will still benefit from the present invention.

Although the different examples have specific components shown in theillustrations, embodiments of this invention are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Although an example embodiment has been disclosed, a worker of ordinaryskill in this art would recognize that certain modifications would comewithin the scope of the claims. For that reason, the following claimsshould be studied to determine their true scope and content.

What is claimed is:
 1. A gas turbine engine comprising: a combustorsection arranged between a compressor section and a turbine section,wherein the compressor section includes at least a low pressurecompressor and a high pressure compressor, the high pressure compressorarranged upstream of the combustor section; a fan section, wherein thelow pressure compressor is downstream from the fan section; and anairfoil arranged in the low pressure compressor and including pressureand suction sides extending in a radial direction from a 0% spanposition to a 100% span position, wherein the airfoil has an axialstacking offset at a 10% span position that is positive and greater thanor equal to an axial stacking offset at the 100% span position, theaxial stacking offsets measured relative to an aftward direction.
 2. Thegas turbine engine according to claim 1, wherein the gas turbine engineis a two-spool configuration.
 3. The gas turbine engine according toclaim 1, wherein the airfoil is rotatable relative to an engine staticstructure.
 4. The gas turbine engine according to claim 1, wherein theairfoil has a relationship between an axial stacking offset and spanposition that includes a curve with a negative slope from 90% to 100%span, where the negative slope leans forward relative to an engine axis.5. The gas turbine engine according to claim 4, wherein the negativeslope is from 80% span to 100% span.
 6. The gas turbine engine accordingto claim 4, wherein the curve has a positive slope from 0% span to 40%span, where the positive slope leans aftward relative to the engineaxis.
 7. The gas turbine engine according to claim 6, wherein the curvetransitions from the positive slope to the negative slope in a range of40% span to 75% span.
 8. The gas turbine engine according to claim 1,wherein the low pressure compressor is counter-rotating relative to thefan section.
 9. The gas turbine engine according to claim 1, wherein thefan section has an array of twenty-six or fewer fan blades and a fanpressure ratio of less than 1.55.
 10. An airfoil for a gas turbineengine comprising: a pressure side and a suction side extending in aradial direction from a 0% span position to a 100% span position; andwherein the airfoil has an axial stacking offset at a 10% span positionthat is positive and greater than or equal to an axial stacking offsetat the 100% span position, the axial stacking offsets measured relativeto an aftward direction.
 11. The airfoil according to claim 10, whereinthe airfoil is rotatable relative to an engine static structure.
 12. Theairfoil according to claim 10, wherein the airfoil has a relationshipbetween an axial stacking offset and span position that includes a curvewith a negative slope from 90% to 100% span, where the negative slopeleans forward relative to an engine axis.
 13. The airfoil according toclaim 12, wherein the negative slope is from 80% span to 100% span. 14.The airfoil according to claim 12, wherein the curve has a positiveslope from 0% span to 40% span, wherein the positive slope leansaftward.
 15. The airfoil according to claim 14, wherein the curvetransitions from the positive slope to the negative slope in a range of40% span to 75% span.
 16. A gas turbine engine comprising: a combustorsection arranged between a compressor section and a turbine section; afan section, wherein the compressor section is downstream from the fansection; a geared architecture coupling the fan section to the turbinesection; and an airfoil arranged in the compressor section and includingpressure and suction sides extending in a radial direction from a 0%span position to a 100% span position, wherein the airfoil has an axialstacking offset at a 10% span position that is positive and greater thanor equal to an axial stacking offset at the 100% span position, theaxial stacking offsets measured relative to an aftward direction. 17.The gas turbine engine according to claim 16, wherein the gearedarchitecture has a gear ratio greater than 2.3.
 18. The gas turbineengine according to claim 16, wherein the airfoil has a relationshipbetween an axial stacking offset and span position that includes a curvewith a negative slope from 90% to 100% span, where the negative slopeleans forward relative to an engine axis.
 19. The gas turbine engineaccording to claim 18, wherein the curve has a positive slope from 0%span to 40% span and transitions from the positive slope to the negativeslope in a range of 40% span to 75% span, wherein the positive slopeleans aftward.
 20. The gas turbine engine according to claim 16, whereinthe fan section has an array of twenty-six or fewer fan blades and a fanpressure ratio of less than 1.55.